Geared architecture turbofan engine thermal management system and method

ABSTRACT

A method of sizing a heat exchanger for a geared architecture gas turbine engine includes sizing a minimum frontal area of at least one heat exchanger located in communication with a fan bypass airflow such that a ratio of waste heat area to horsepower generation characteristic area is between 1.6 to 8.75.

CROSS-REFERENCE TO RELATED APPLICATIONS

The instant application is a continuation application of U.S. patentapplication Ser. No. 14/776,779, filed Sep. 15, 2015, which is a 371 ofInternational Application PCT/US2014/025552, filed Mar. 13, 2014, whichclaims the benefit of provisional application Ser. No. 61/792,395, filedMar. 15, 2013.

BACKGROUND

The present disclosure relates to a geared architecture gas turbineengine and, more particularly, to a heat exchanger therefor.

Gas turbine engines, such as those which power modern commercial andmilitary aircraft, include a compressor section to pressurize a supplyof air, a combustor section to burn a hydrocarbon fuel in the presenceof the pressurized air, and a turbine section to extract energy from theresultant combustion gases and generate thrust.

Aero engine Thermal Management Systems (TMS) typically include heatexchangers and associated equipment which exchange engine heat with anairflow or fuel flow. The gas turbine engine architecture typicallydictates TMS heat exchanger placement.

SUMMARY

A method of sizing a heat exchanger for a geared architecture gasturbine engine according to one disclosed non-limiting embodiment of thepresent disclosure includes sizing a minimum frontal area of at leastone heat exchanger located in communication with a fan bypass airflowsuch that a ratio of waste heat area to horsepower generationcharacteristic area is between 1.6 to 17.5.

A further embodiment of the present disclosure includes, wherein thewaste heat area is defined by the minimum frontal area of the HEX.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the horsepower generation characteristicarea is defined by an exit area of a high pressure compressor.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, locating the at least one heat exchanger within afan bypass airflow path such that the ratio of waste heat area tohorsepower generation characteristic area is between 1.6 to 8.75.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, locating the at least one heat exchanger withrespect to a fan duct total pressure profile.

A method of sizing a heat exchanger for a geared architecture gasturbine engine according to another disclosed non-limiting embodiment ofthe present disclosure includes determining an efficiency of a gearedarchitecture; determination a temperature requirement of the oil at aparticular flight condition; determining a fan pressure ratio; andsizing a minimum frontal area of the at least one heat exchanger inresponse to the efficiency of the geared architecture, the temperaturerequirements of the oil at a particular flight condition, and the fanpressure ratio.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, sizing the minimum frontal area of the at least oneheat exchanger such that a ratio of waste heat area to horsepowergeneration characteristic area is between 1.6 to 17.5.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the geared architecture provides anefficiency of at least 97.7%.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the fan pressure ratio is less than 1.5.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the at least one heat exchanger provides anefficiency of at least 50%.

A gas turbine engine according to another disclosed non-limitingembodiment of the present disclosure includes a geared architecture thatprovides an efficiency above 97.7%; a fan driven by the gearedarchitecture to generate a fan bypass airflow; and at least one heatexchanger mounted in communication with the fan bypass airflow, aminimum frontal area of the at least one heat exchanger defines an arealess than 420 in{circumflex over ( )}2 (270967 mm{circumflex over( )}2).

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the geared architecture provides anefficiency of at least 97.7%.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the fan provides a fan pressure ratio ofless than 1.5.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the at least one heat exchanger provides acooling efficiency of at least 50%.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the minimum frontal area of the at leastone heat exchanger is sized such that a ratio of waste heat area tohorsepower generation characteristic area is between 1.6 to 17.5.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the geared architecture provides anefficiency of at least 97.7% and the at least one heat exchanger anprovides efficiency of at least 50%.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the fan provides a fan pressure ratio ofless than 1.5 and the at least one heat exchanger provides an efficiencyof at least 50%.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the geared architecture provides anefficiency of at least 97.7%, the fan provides a fan pressure ratio ofless than 1.5, and the at least one heat exchanger provides anefficiency of at least 50%.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the fan bypass airflow provides a bypassratio greater than 6.0.

A further embodiment of any of the foregoing embodiments of the presentdisclosure includes, wherein the at least one heat exchanger is operableto maintain an oil temperature below 325 F (163 C).

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of a gas turbine engine;

FIG. 2 is an expanded schematic cross-section of a fan airflow path ofthe gas turbine engine;

FIG. 3 is graphical representation of HEX sizing relationships for ageared architecture gas turbine engine;

FIG. 4 is schematic representation of T3 aft of the last annular flowarea in the plane of the trailing edge of the last blade annular flowarea in the plane of the trailing edge of the last blade of the HPC; and

FIG. 5 is a flow chart illustrating sizing boundaries of a heatexchanger in communication with the fan bypass airflow for a gearedarchitecture gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a fan bypass flowpathwhile the compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be appreciated that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines such as a three-spool (plus fan) engine wherein anintermediate spool includes an intermediate pressure compressor (IPC)between the LPC and HPC and an intermediate pressure turbine (IPT)between the HPT and LPT.

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine case assembly 36 via several bearing structures38. The low spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 (“LPC”) and a lowpressure turbine 46 (“LPT”). The inner shaft 40 drives the fan 42through a geared architecture 48 to drive the fan 42 at a lower speedthan the low spool 30.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”). Acombustor 56 is arranged between the high pressure compressor 52 and thehigh pressure turbine 54. The inner shaft 40 and the outer shaft 50 areconcentric and rotate about the engine central longitudinal axis A whichis collinear with their longitudinal axes.

Core airflow is compressed by the low pressure compressor 44 then thehigh pressure compressor 52, mixed with the fuel and burned in thecombustor 56, then expanded over the high pressure turbine 54 and thelow pressure turbine 46. The turbines 54, 46 rotationally drive therespective low spool 30 and high spool 32 in response to the expansion.

In one non-limiting embodiment, the gas turbine engine 20 is ahigh-bypass geared architecture engine in which the bypass ratio isgreater than six (6:1). The geared architecture 48 can include anepicyclic gear train, such as a planetary gear system, star gear systemor other gear system. The example epicyclic gear train has a gearreduction ratio of greater than 2.3, and in another example is greaterthan 2.5. The geared turbofan enables operation of the low spool 30 athigher speeds which can increase the operational efficiency of the lowpressure compressor 44 and low pressure turbine 46 and render increasedpressure in a fewer number of stages.

A pressure ratio associated with the low pressure turbine 46 is pressuremeasured prior to the inlet of the low pressure turbine 46 as related tothe pressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle of the gas turbine engine 20. In one non-limitingembodiment, the bypass ratio of the gas turbine engine 20 is greaterthan ten (10:1), the fan diameter is significantly larger than that ofthe low pressure compressor 44, and the low pressure turbine 46 has apressure ratio that is greater than five (5:1). It should beappreciated, however, that the above parameters are only exemplary ofone embodiment of a geared architecture engine and that the presentdisclosure is applicable to other gas turbine engines including directdrive turbofans.

The high bypass ratio results in a significant amount of thrust. The fansection 22 of the gas turbine engine 20 is designed for a particularflight condition—typically cruise at 0.8 Mach and 35,000 feet. Thisflight condition, with the gas turbine engine 20 at its best fuelconsumption, is also known as bucket cruise Thrust Specific FuelConsumption (TSFC). TSFC is an industry standard parameter of fuelconsumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of (“T”/518.7)^(0.5). The Low Corrected Fan Tip Speedaccording to one non-limiting embodiment of the example gas turbineengine 20 is less than 1150 fps (351 m/s). The Low Corrected Fan TipSpeed in another non-limiting embodiment of the example gas turbineengine 20 is less than 1200 fps (366 m/s).

With reference to FIG. 2, the fan 42 drives air along a fan bypassflowpath W past a Fan Exit Guide Vane system 60. A thermal managementsystem (TMS) 62 includes a heat exchanger (HEX) 64 that may be at leastpartially integrated into a nacelle assembly 66 such as a fan nacelle68, a core nacelle 70, a bifurcation 72, the Fan Exit Guide Vane system60 or any combination thereof but still considered as in communicationwith the fan bypass flowpath W as defined herein.

The HEX 64 in the disclosed non-limiting embodiment may be a “brick”type HEX 64 that, for example, may include an air-oil cooler or coolerarray that services both the geared architecture 48 and/or engine oilcircuits such as that which communicates with the bearing compartments.The HEX 64 in another disclosed non-limiting embodiment may beselectively moved into the fan bypass flowpath W when oil temperature orfuel temperature are cooler and less HEX frontal area and fan pressureloss are desired. The HEX 64 in still another disclosed non-limitingembodiment may be mounted within the nacelle assembly 66, the HEX 64being positioned downstream of a scooped duct 100. Further, it should beappreciated that the minimum frontal area 65 of the HEX 64 may bedistributed over a multiple of heat exchangers and that the examplesizes provided herein are only utilized to depict relative sizing of thefrontal area 65 of the HEX 64 and are not to be considered limiting.

Generally, HEX 64 location [A] is defined herein as the center or“heart” of the fan duct total pressure profile F across the fan bypassflowpath W while the HEX 64 location [B] is outboard of location [A] andthe HEX 64 location [C] is located inboard of location [A].Alternatively, the HEX 64 location [D] is a buried arrangement, hereshown as within the core nacelle 70 with a scooped duct system 100 thatcommunicates bypass airflow thereto. The HEX 64 at location [A] may besmaller than at location [B], and HEX 64 at location [B] may be smallerthan a HEX 64 at location [C] because the total pressure from the fan 42is greatest at a radial central location [A] and lowest at location [C].That is, the fan duct total pressure profile F varies radially acrossthe fan bypass flowpath B. The total pressure from the fan 42 and thusthe HEX 64 efficiency varies at least in part by the efficiency of thegeared architecture 48 and the Fan Pressure Ratio.

In addition, the efficiency of HEX 64 location [D] is also effected bylocation with respect to the scooped duct 100. It should be understoodthat other factors may also contribute to HEX efficiency, i.e., airtemperature rise vs. maximum air temperature rise of the HEX—for exampleif oil is at a temperature of 300 degrees Fahrenheit (300 F) and fanairflow is at a temperature of 100 F (38 C), a 100% effective HEX wouldreduce the oil temperature from 300 F (149 C) to 100 F (38 C). That is,for example, fan duct pressure loss, system-level optimization of heatexchanger weight, the effect of fan duct losses on fuel burn, and othersystem-level effects may also be accounted for as a product ofcomputerized numerical calculations and empirical data on, for example,a machine readable storage medium having stored thereon a computerprogram for sizing of the minimum frontal area 65 of the HEX 64 for ageared architecture gas turbine engine.

With reference to FIG. 3, the position of the HEX 64 [A, B, or C; FIG.2]; the efficiency of the HEX 64 [90% or 50%]; the Fan Pressure Ratio(FPR) [_(1.4; 1.3; or 1.2)] and the efficiency of the gearedarchitecture 48 [97.7%; 98.7%; or 99.7%] variables are plotted withrespect to: a flat rated takeoff thrust takeoff at 86° F. (30° C.); anannular exit area of the final rotor stage in the HPC 52 (illustrated inFIG. 4) [X-axis]; and a HPC 52 exit corrected flow with respect to thefrontal area 65 of the HEX 64 [Y-axis]. It should be appreciated thatother relationships may alternatively be defined.

The fan pressure ratio (FPR) is representative of the total pressureavailable to the HEX 64. Only the minimum frontal area 65 of the HEX 64need be considered and this minimum frontal area 65 of the HEX 64 may benormalized and given with respect to an annular flow area in the planeof the trailing edge of the final rotor stage 52-1 (FIG. 1) in the HPC52 (FIG. 4) which is representative of the overall size of the engine20. It should be appreciated that other relationships may be providedwhich result in equivalent HEX 64 sizing.

Positions A90 _(1.4; 1.3; 1.2) locates the HEX 64 at mid-stream of thefan airflow W with an example 90% HEX 64 efficiency where air exittemperature measured anywhere at the exit of the HEX 64 is approximately90% of the oil input temperature. This is a relatively large HEX 64.

Positions B90 _(1.4; 1.3; 1.2) locate the HEX 64 at the outer diameterof the fan bypass flowpath W. Position C90 locates the HEX 64 at theinner diameter of the fan bypass flowpath W. Positions A50_(1.4; 1.3; 1.2) sacrifice HEX 64 air temperature; for example, theefficiency is reduced to 50%, at the respective three locations A, B, Cbut the HEX is relatively smaller. At positions C50 _(1.4; 1.3; 1.2) thefrontal area 65 of the HEX 64 is now much larger than positions A90_(1.4; 1.3; 1.2); which describes the maximum and minimum frontal area65 of the HEX 64 for an un-ducted HEX 64. At positions C50_(1.4; 1.3; 1.2) the HEX 64 is relatively large because the fan pressureis extremely low and the air temperature rise is only 50% of the idealtemperature rise. Thus the minimum frontal area 65 of the HEX 64 C50_(1.4; 1.3; 1.2) to the maximum frontal area of the HEX 64 at positionsA90 _(1.4; 1.3; 1.2) is enveloped between these parameters.

The area between positions A90 _(1.4;) and C50 _(1.2;) is furtheraffected by the efficiency of the geared architecture 48. A boundary Fis when the geared architecture 48 efficiency is less than or equal to97.7%; the fan pressure ratio (FPR) is, e.g., 1.4 (although it could behigher such as 1.50); and the HEX 64 is efficient e.g., provides 90%efficiency. It should be appreciated that all three exampleparallelograms have the same constituent components for positions A, Band C but are not all drawn here for clarity. That is, the mostefficient 99.7% geared architecture is shown combined with the Apositions while the least efficient 97.7% geared architecture is showncombined with the C positions to provide the greatest bounds.Furthermore, position A90 _(1.4) with 99.7% efficiency while positionC50 _(1.2) with 97.7% efficiency provide an example relative practicalrange of minimum frontal area 65 of the HEX 64 as sized with respect tothe fan bypass airflow. The disclosed relationship facilitatesdetermination of minimum frontal area 65 of the HEX 64 for a gearedarchitecture gas turbine engine. For example, 24,000 pounds of thrust atSea Level Takeoff, flat rated to an 86° F. day, with a 97.7% efficiencygeared architecture 48 may be plotted so that the HEX 64 may then besized to a particular altitude and fuel temperature condition.

With reference to FIGS. 3 and 5, initially, one boundary condition tosize the minimum frontal area 65 (of at least one HEX 64 mounted incommunication with a fan bypass airflow) is determining the efficiencyof the geared architecture 48, here shown as three general examples ofgeared architecture efficiency: high, e.g., the 99.7% “A” plane; medium,e.g., the 98.7% “B” plane; and low, e.g., the 97.7% “C” plane. Thisefficiency of the geared architecture 48 is herein described as thestarting point and primary variable that will affect the minimum frontalarea 65 required as shown on the Y-axis of FIG. 3.

Another boundary condition is the determination of the temperaturerequirement of the oil, for example, a maximum oil temperature of 325°F. (163° C.) leaving the engine. Accordingly, for the same efficiency ofthe geared architecture 48, a higher allowable temperature will resultin a relatively smaller minimum frontal area 65 of the HEX 64 while alower allowable temperature will result in a relatively larger minimumfrontal area 65. It should be appreciated that other factors such asother internal fuel-oil heat exchangers may be involved and areaccounted for in the primary variable discussed, i.e, the maximum oiltemperature is the resultant temperature through the system. Thus, allof the factors in this paragraph are depicted in FIG. 3, as the areabetween positions A90 _(1.4) and C50 _(1.2).

Another boundary condition involved in the design of the engine is thefan pressure ratio at the engine thrust required by the aircraft.Realistically, as the designer decreases engine fan pressure ratio toimprove fuel burn, the engine bypass ratio increases until the engineweight and drag increase such that the fan duct size causes the fuelburn to reach a minimum. The fan pressure ratio is that which drives themass flow into the HEX 64 so that as fan pressure drops, the minimumfrontal area 65 of the HEX 64 will be relatively larger. This isdepicted in FIG. 3 as the upward sloping lines of increasing area withincreasing thrust but also increasing area at the same thrust if one iscomparing a 1.4 fan pressure to a 1.2 fan pressure.

An example calculation is based around an engine with 33,000 lbs ofstatic thrust at sea level on an 86° F. day and a 1.36 fan pressureratio at 35K feet altitude. Takeoff operations provide significant fuelflow cooling of the oil via a fuel-oil HEX (not shown), so takeoffoperations are not the air-oil HEX 64 sizing point. The minimum sizingpoint is for flight operations, which, in this example, are with 17,424pounds of thrust. Notably, thrust decreases at altitude due to reducedair density. At this example sizing condition, the 99.7% efficiencygeared architecture 48, with a target heat exchanger efficiency of 72%will have a minimum frontal area 65 of 45 in{circumflex over ( )}2(29032 mm{circumflex over ( )}2) to arrive at an oil temperature maximumof 325° F. maximum. For the same example, a 98.7% efficiency gearedarchitecture 48, would require a minimum frontal area 65 of 130in{circumflex over ( )}2 (83870 mm{circumflex over ( )}2) while a 97.7%efficiency geared architecture 48 would require a frontal area 65 of 210in{circumflex over ( )}2 (135484 mm{circumflex over ( )}2) to providethe desired maximum oil temperature of 325° F. A buried architectureminimum frontal area 65, e.g., location [D] (FIG. 2) would be 420in{circumflex over ( )}2 (270967 mm{circumflex over ( )}2).

The combination of these and other variables may also be satisfactorilynormalized and compared to an exit area S of the final rotor stage 52-1(FIG. 1) in the HPC 52 (FIG. 4) as all the airflow that is mixed with,and ultimately exits this engine section may be considered the presentstate of the art for maximum combustor exit temperature that is viablefor acceptable engine durability and economics. That is, the exit area Sis a valuable figure of merit to characterize the horsepower that can beproduced by the engine 20 and the minimum frontal area 65 of the HEX 64can be compared thereto, since it represents the handling of lostenergy.

For one example, the minimum frontal area 65 range of the HEX 64required to cool the oil to a desired temperature can range from 45-420in{circumflex over ( )}2 as described above and the exit area S of thefinal rotor stage 52-1 is 28 in{circumflex over ( )}2. In one disclosednon-limiting embodiment, the ratio of “waste heat area”, i.e., thefrontal area 65 to “horsepower generation characteristic area” i.e., theexit area S of the last annular flow area in the plane of the trailingedge of the final rotor stage 52-1 ranges from 1.6 to 8.75 and isrelatively linear for all engines using air-oil heat exchangers that seefan total pressure directly as illustrated in locations [A], [3], [C] ofFIG. 2. In another disclosed non-limiting embodiment, heat exchangersthat are buried inside of parasitic ducts adjacent to the fan bypassflowpath W with scoops within the fan bypass flowpath W that capture fantotal pressure, e.g., location [D], duct losses, valves etc., result ina “waste heat area” to “horsepower generation characteristic area” of3.2 to 17.5. That is, the buried HEX 64 arrangement would have a “wasteheat area” to “horsepower generation characteristic area” that isessentially double that of a heat exchanger location [A] within the fanbypass flowpath W, i.e., 420 in{circumflex over ( )}2.

It should be appreciated that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be appreciated that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be appreciated that steps may be performed in any order,separated or combined unless otherwise indicated and will still benefitfrom the present disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beappreciated that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A method of sizing a heat exchanger for a gearedarchitecture gas turbine engine comprising: sizing a minimum frontalarea of at least one heat exchanger located in communication with a fanbypass airflow of the geared architecture gas turbine engine downstreamof a fan such that a ratio of waste heat area to horsepower generationcharacteristic area is between 1.6 to 17.5, the waste heat area definedby the frontal area of the at least one heat exchanger, wherein thehorsepower generation characteristic area is defined by an exit area ofa high pressure compressor of the geared architecture gas turbineengine.
 2. The method as recited in claim 1, further comprising locatingthe at least one heat exchanger within a fan bypass airflow path.
 3. Themethod as recited in claim 1, further comprising locating the at leastone heat exchanger within a fan bypass airflow path such that the ratioof waste heat area to horsepower generation characteristic area isbetween 1.6 to 8.75.
 4. The method as recited in claim 1, furthercomprising locating the at least one heat exchanger with respect to afan duct total pressure profile.
 5. A geared architecture gas turbineengine comprising: a geared architecture; a fan driven by the gearedarchitecture of the gas turbine engine to generate a fan bypass airflow;and at least one heat exchanger mounted in communication with the fanbypass airflow downstream of the fan, a minimum frontal area of the atleast one heat exchanger defines an area less than 420 in² (270967 mm²)and greater than 0 in², the at least one heat exchanger comprising aratio of waste heat area to horsepower generation characteristic area isbetween 1.6 to 17.5, the waste heat area defined by the frontal area ofthe at least one heat exchanger, wherein the horsepower generationcharacteristic area is defined by an exit area of a high pressurecompressor of the geared architecture gas turbine engine.
 6. The gearedarchitecture gas turbine engine as recited in claim 5, wherein the fanprovides a fan pressure ratio of less than 1.5.
 7. The gearedarchitecture gas turbine engine as recited in claim 5, wherein the atleast one heat exchanger provides a cooling efficiency of at least 50%.8. The geared architecture gas turbine engine as recited in claim 5,wherein the geared architecture provides an efficiency of at least 97.7%and the at least one heat exchanger provides an efficiency of at least50%.
 9. The geared architecture gas turbine engine as recited in claim5, wherein the fan bypass airflow provides a bypass ratio greater than6.0.
 10. The geared architecture gas turbine engine as recited in claim5, wherein the at least one heat exchanger is operable to maintain anoil temperature below 325 F (163 C).
 11. The geared architecture gasturbine engine as recited in claim 5, wherein the at least one heatexchanger is located at an outer diameter of the fan bypass airflow. 12.The geared architecture gas turbine engine as recited in claim 5,wherein the at least one heat exchanger is located at mid-stream of thefan bypass airflow.
 13. The geared architecture gas turbine engine asrecited in claim 5, wherein the at least one heat exchanger is locatedat an inner diameter of the fan bypass airflow.
 14. A gearedarchitecture gas turbine engine comprising: a geared architecture; a fandriven by the geared architecture of the gas turbine engine to generatea fan bypass airflow; and at least one heat exchanger mounted incommunication with the fan bypass airflow downstream of the fan, the atleast one heat exchanger comprising a ratio of waste heat area tohorsepower generation characteristic area is between 1.6 to 17.5, thewaste heat area defined by the frontal area of the at least one heatexchanger, wherein the horsepower generation characteristic area isdefined by an exit area of a high pressure compressor of the gearedarchitecture gas turbine engine.